Compressor blading



April 26, 1960 G. F. HAusMANN COMPRESSOR BLADING 2 Sheets-Sheet i FiledJune 18, 1956 Tf/e r y t nx.

/ ,ear/swam cam/E/vr/o/VAL ma` mp/4f@ Fo@ Max/MUM N VE N TOR GEORGE EHUSMNN Br vf/ A T TORNEY April 26, 1960 G. F. HAUSMANN 2,934,259

COMPRESSOR BLADING Filed June 18, 1956 2 Sheets-Sheet 2 M (HX/HL)Max/Mam rH/cA/A/ss /N VEN TOR GEORGE E HAUSMANN ATTORNEY United 'StatesPatent O F COMPRESSOR BLADING George F. Hausmann, Glastonbury, Conn.,assignor to United Aircraft Corporation, East Hartford, Conn., acorporation of Delaware Application June 18, 1956, Serial No. '592,118

4 Claims. (Cl. 5230-122) This invention relates to compressors and morespecically to axial flow transonic or supersonic compressors.

It is an object of this invention to provide a compressor having a bladeconiiguration which provides maximum ecieney for a given pressure riseacross the blading while conforming to structural limitations imposed bystress considerations.

This and other objects of this invention will become readily apparentfrom the following detailed description of the drawings in which:

Fig. 1 is a partial schematic and partial cross-sectional View of acompressor rotor or stator having conventional blading therein.

Fig. 2 is similar to Fig. 1 but illustrates the proposed blading of thisinvention.

Fig. 3 is a schematic illustration in partial cross section showing amodified form of blading according to this invention, and

Fig. 4 is a plan view of one of the blades of Fig. 3.

1t has been determined from two-dimensional transonic and supersoniccascade tests that the total pressure losses are minimized in axial owcompressors for most conventional blade shapes when the blades arearranged at a gap-chord ratio (r/c) of approximately .5. However, theuse of this gap-chord ratio (f/c) at the tip sections (when the relativevelocities are transonic or supersonic) requires a gap-chord ratio atthe root which is lower in proportion to the hub-tip diameter ratio.Since the use of a gap-chord ratio lower than .3 or .4 at the root ofthe blades of an axial flow compressor would severely compromise theperformance of these sections (choking or low critical Mach numbers), itwould be necessary to utilize inverse taper on the planform of the bladeto obtain the desirable aerodynamic characteristics at the root and tip.Furthermore, it would also complicate the mechanical structure of theblade attachment fixtures because of the space limitations on thecompressor rotor disc. However, the use of inverse taper imposes a lowstress limit on the blade.

The blade passage described herein would provide the desirableaerodynamic characteristics of a low gap-chord ratio at the compressortip without resorting to inverse planform taper.

Fig. l depicts an array of blades at the tip section of a conventionaltransonic or supersonic axial flow compressor. '-It has been determinedfrom experiment that blade passages of this type have a minimum totalpressure loss for a given pressure rise and turning, which describe thepressure rise in the corresponding compressor element. This particularcascade has a gap-chord ratio (T/c) of approximately .5, which isundesirable from compressor stress considerations. The maximum thicknessof the blades also occurs at approximately the 50% chord position. Thefavorable aerodynamic performance of this cascade is due to thecharacteristics of the flow at the location of the intercept of theshock wave from the leading edge of the succeeding blade on the uppersurface of the preceding blade. The Mach Y P2,934,259 Patented Apr. 26,1960 ICC number at this intercept is essentially the same as therelative inlet Mach number and the boundary layer thickness is lowerthan that at any other -point within the passage. If the gap-chord ratio(r/c) were increased, the Mach number and boundary layer thickness atthe shock intercept would both be increased. Downstream of the shockintercept in Fig. 1, the ow diffuses and turns subsonically. However,the pressure rise and compressor work associated with this process aresmall relative to the turning and pressure rise which is obtained acrossthe shock wave system.

In order to obtain these desirable aerodynamic characteristics withoutseverely compromising the structural considerations, it is proposed toterminate the chord of the blades shown in Fig. l at a position slightlydownstream of the shock intercept, as shown in Fig. 2. This modicationplaces the maximum thickness at approximately the -percent chord station'and results in a blunt, or relatively blunt trailing edge. In thisrespect, the subsonic diffusion and turning would be compromised.However, it has been demonstratedby test that this cornpromise is notsevere. The test results have indicated that the favorableshock-interaction characteristics are maintained, whereas the gap-chordratio is essentially doubled.

It is believed that the nomenclature shown in Figs. 1 and 2 issufficiently clear so that extended application of reference charactersis not necessary. lt might be pointed out, however, that in the Fig. 2configuration according to this invention the maximum thickness of theblading is approximately at the location indicated by the numeral 10.This location may range from 75% of the chord, as shown, or in anyposition aft thereof to the chord location. This makes the gap-chordratio (r/c) ranging between .75 to 1.0. It should be added that the lowpressure side of each blade will have a surface forward of the maximumthickness point which is substanitally aligned or`parallel to thedirection of the relative inlet flow. The point of maximum thicknessthen forms the passage throat between the blades in cooperation with thebottom surface (high pressure side) of the next adjacent blade. Thebottom surface of each blade will be at an angle to the relative inletow an amount approximately equal to the Wedge angle of the blade.

The blade section of this invention would be used only on the outboardportion of the blades where the relative approach Mach numbers aregreater than one. Thus the type of blade shown in Figs. 3 and 4 is oneembodiment of this invention. In other words, the flow becomes morecritical toward the outboard end of the blades 20, 22 so that themaximum thickness may be located as, for example, as shown respectivelyat 24 and 26 in the 50% chordwise location at the root portion. However,adjacent the tip portion of these blades the maximum thickness point 30is located in the last quarter of the chordwise dimension of the bladesection at that point. The blade 20 is better illustrated in Fig. 4illustrating that the maximum thickness point of the blade 24 is locatedapproximately at the 50% chordwise position at the root portion of theblade and the maximum thickness point changes until it reachesapproximately a location within the last 25% of the chordwise dimensionadjacent the tip of the blade 20.

it should be added that the trailing edge of the blades of thisinvention aft of the maximum thickness point is relatively bluntcompared to the leading edge of the blades. In other words, aft of themaximum thickness point the blade may be made relatively blunt orrounded to any suitable shape commensurate with good design practice andfabrication limitations.

As a result of this invention, it is apparent that a sim-` ple yethighly efficient blade configuration has been prothe construction andarrangement of the various parts n without departing from the scope ofthis novel concept.

I claim:

1. A compressor comprising a casing structure, a compressor rotorjournaled axially in said casing structure and defining therewith anannular ow passage, blading carried by said rotor and extending acrosssaid annular flow passage in a cascade arrangement, said rotor bladesbeing substantially arfoil in section from root to tip and having asubstantially sharp leading edge and a relatively blunt trailing edge,said blades having their point of maximum thickness adjacent saidtrailing edges, the low pressure or top of the leading blade having atleast two surfaces at an angle relative to each other and the highpressure or bottom of the next adjacent blade having a single surface ata small angle relative to the oncoming stream, said blades ybeingarranged so that the inlet edge of each blade together with the maximumthickness portion of the adjacent leading blade defines a restrictedthroat in the passage between the blades, said throat beingsubstantially aligned with the direction of how of the gaseous mediumflowing through the blade passage for effecting a normal compressionshock to gaseous medium entering said throat at superacoustic velocityrelative to the moving blades and thereby provide an immediate reductionin the velocity of said medium and an immediate increase in the pressurethereof, the iiow downstream of said shock being subsonic and confinedin a passage defined by only one of two of said adjacent blades of saidcascade.

2. A compressor comprising a casing structure, a compressor rotorjournaled axially in said casing structure and defining therewith anannular fiow passage, blading carried by said rotor and extending acrosssaid annular flow passage in a cascade arrangement, stator vanes carriedby said casing structure and extending across said flow passagedownstream of said rotor blades and arranged to direct the gaseousmedium to be compressed at a proper angle to the next adjacent rotor,said rotor blades being substantially airfoil in section from root totip and havingr a substantially sharp leading edge and relatively blunttrailing edge, said blades having their point of maximum thickness inthe 75% to 100% chordwise position, said blades being arranged so thatthe inlet edge of each blade together with the maximum thickness portionof the adjacent leading blade delines in cooperation with the leadingedge of the next following blade a restricted throat in the passagebetween the blades, said throat being substantially aligned with thedirection of flow of the gaseous medium iiowing through the bladepassage for effecting a normal compression shock to gaseous mediumentering said throat at superacoustic velocity relative to the movingblades and thereby provide an immediate reduction in the velocity ofsaid medium and an immediate increase in the pressure thereof.

3. A compressor comprising a casing structure, a compresser rotorjournaled axially in said casing structure and defining therewith anannular flow passage, blading carried by said rotor and Vvextendingacross said annular flow passage in a cascade arrangement, said rotorblades being substantially airfoil in section from root to tip andhaving a substantially sharp leading edge and a relatively blunttrailing edge, said blades having top and' bottom surfaces which definethe low and high pressure sides of each of said blades, respectively,said blades having their point of maximum thickness adjacent saidtrailing edges and said point being formed on the top or low pressurecambered surface of each blade, said point being defined by a maximumdeparture of said top or low pressure surface from the chord line of theblade, said blades being arranged so that the inlet edge of each bladetogether with the maximum thickness portion of the adjacent leadingblade defines a restricted throat in the passage between the blades,said top surface upstream of said throat being substantially alignedwith the relative inlet flow and said throat being substantially alignedwith the direction of iiow of the gaseous medium owing through the bladepassage for effecting a normal compression shock to gaseous mediumentering said throat at superacoustic velocity relative to the movingblades and thereby provide an immediate reduction in the velocity ofsaid medium and an immediate increase in the pressure thereof, the flowdownstream of said shock being subsonic and confined in a passagedefined by only one of two of said adjacent blades of said cascade.

4. In an axial flow supersonic compressor having at least one rotor,said rotor including a plurality of circumferentially spaced bladeshaving a predetermined angle of attack relative to the axis of rotationof said rotor, said blades in cross section having relatively lsharpleading edge portions extending a substantial portion of the chord- Wisedimension of the blades, said blades having top or low pressure andbottom or high pressure surfaces with the top surface leading in motionduring rotation of the blades about said axis, the low pressure or topsurface of said blades having upstream and downstream surface por tionswith the upstream portion extending over a majority of the chordwiselength ofthe blade and being substantially parallel to the relativedirection of the inlet flow, the downstream surface portion being at anangle relative to said upstream portion, and the point of maximumthickness of the blades adjacent the outer tip thereof being locatedadjacent the trailing edge of said blades thereby forming a throat withthe leading edge of the following blade whereby the leading edge shockextending at substantially right angle from the high pressure surface ofone blade intercepts the trailing edge region of the low pressuresurface of the next adjacent blade approximately inthe to 100% chordwiserange.

References Cited in the file of this patent UNITED STATES PATENTS2,721,693 Fabri et al. Oct.r25, 1955

